Method of making turbine blades and the like



June 13, 1961 E. G. PEKAREK 2,987,806

D OF MAKING TURBINE BLADES AND THE LIKE Filed May 24, 1956 2 Sheets-Sheet 1 V5.27 [57F WARD G. PE/(Afiii r by y H/gs.

June 13, 1961 E. G. PEKAREK 2,987,806

METHOD OF MAKING TURBINE BLADES AND THE LIKE Filed May 24, 1956 2 Sheets-Sheet 2 00 z 2 in s D B: O u n:

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[HZ/E27 [UP United States Patent 2,987,806 METHOD OF MAKING TURBINE BLADES AND THE LIKE Edward G. Pekarek, Willoughby, Ohio, assignor to gilpmpson Ramo Wooldridge Inc., a corporation of Filed May 24, 1956, Ser. No. 586,996 2 Claims. (Cl. 29-1563) The present invention is concerned with a method for making articles such as turbine blades and the like.

The manufacture of parts such as turbine buckets, compressor blades, nozzle diaphragm vanes, and the like, presents the problem of providing a complex shape to within very close tolerances. Coupled with this requirement is the necessity of employing a strong, creep resistant metal or alloy for the body of the part. In use, turbine blades must be capable of withstanding a combination of complex, superimposed thermal fatigue and bending stresses so that a suitable metal or alloy must have a rather unique combination of properties to be suitable. In some cases, metals or alloys which are sufficiently strong, hard, and have good oxidation resistance cannot be adapted readily, due to their refractory nature, to conventional forging practices without considerable difficulty. As a result, such metals or alloys require the employment of rather complex processes for their fabrication. Even with such processes, however, the number of rejects is frequently high and the processes are necessarily expensive.

The present invention overcomes these difficulties by providing a process for the manufacture of turbine blades and the like consisting essentially of a casting step followed by a hot working step in which the cast article is brought to its final dimensions and in which selected portions of the turbine blade are simultaneously provided with the proper metallurgical properties in order to resist adequately the specific types of stresses which those portions will undergo in operation. For example, a turbine bucket is subjected during use to various types of stresses in different portions of the bucket. The root area of the bucket is subject to creep stresses due to the effect of centrifugal force. In contrast, the tip end of the vane, and the leading and trailing edges of the turbine bucket vane portion are subjected to fatigue stresses resulting from vibration and thermal stresses due to rapid changes in temperature along the edges. The differences in the nature of the stresses occurring in difierent parts of the turbine blade require, for the longest operating life, that the different portions have different physical properties. In the process of the present invention, these different physical properties are achieved by first casting the metal into an oversized preform of the turbine blade, with excess metal being provided at the areas in which resistance to fatigue is to be achieved, and following the casting by a hot workin operation which simultaneously reduces the dimensions of the article in the oversized areas down to the finished dimensions, and changes the as cast structure of the metal into a form which is better adapted to resist fatigue stresses.

The casting of the metal or alloy initially is preferably done under vacuum conditions or at least by casting a master melt which has previously been melted under vacuum in the presence of an inert atmosphere such as argon, helium, or the like. Some of the improved results obtained through the process of the present invention have been achieved where the initial casting was done in the presence of air, but vacuum melted material yields considerably better results.

An object of the present invention is to provide an improved process for the manufacture of turbine blades and the like which provide selected areas on the turbine blade with metallurgical properties best suited to resist the particular stresses which that portion will undergo in use.

Another object of the present invention is to provide a process for the manufacture of turbine blades and the like from alloys which are difiicult to forge from bar stock or ingot form.

Another object of the present invention is to provide a process for the manufacture of turbine blades and the like which makes it highly convenient to repair surface defects in the article during the operation of the process.

The process of the present invention may employ nickel or cobalt alloys which have heretofore been of limited utility in the manufacture of turbine blading because of difi'iculties arising from forging such alloys from bar stock or ingot form. Particularly good results have been obtained with nickel base, precipitation hardenable alloys containing from about 13 to 21% chromium, from 1.0 to 6.0% molybdenum, from 1.0 to 4.0% aluminum, from 1.5 to 3.5% titanium, a maximum of about 0.2% carbon, and in some cases, from about 10 to 35% cobalt, with the balance being substantially nickel. A particularly preferred alloy has the following composition:

Table I C 0.1 to 0.2% by weight. Mn 0.25% maximum. Si 0.60% maximum. Cr 14 to 17%. Mo 4.5 to 6.0%. Ti 1.5 to 2.5%. A1 2.5 to 3.5%. B .025 to .07%. Fe 8 to 12%. Ni Substantially the balance.

Another suitable alloy is that known commercially as Waspaloy which has the following analysis:

Table II C 0.1% maximum. Cr 18.0 to 21.0%. C0 12.0 to 15.0%. Mo 3.5 to 5.0%. A1 1.0to 1.5%. Ti 2.75 to 3.25%. Ni Substantially the balance.

Still another nickel base alloy which has been successfully employed is known as Inco-700" and has the following analysis:

Table III C 0.16% maximum. Cr 13.0 to 17.0%. C0 24.0 to 34.0%. Mo 1.0 to 4.5%. A1 2.5 to 3.5%. Ti 1.75 to 2.75%. Ni Substantially the balance.

The original casting is made under vacuum conditions in order to secure a higher degree of ductility in the casting. For making the casting, pressures on the order of less than one millimeter of mercury, and preferably not in excess of about microns of mercury are appropriate. Satisfactory results can also be obtained by remelting a master melt previously made under such vacuum conditions in an inert atmosphere.

The casting which results, when cooled relatively slowly, as in conventional investment casting practice usually exhibits a coarse crystalline structure characterized by the presence of dendrites. One of the features of the e 3 process of the present invention resides in maintaining as cast structure in portions of the turbine blade, particularly in the relatively massive root portion of the blade andin the airfoil area adjacent to the root. Testshave indicated that the as cast structure provides a. better resistance to creep stresses than does a relatively fine crystalline structure typical of a hot worked alloy composition, While the fine grained material is inherently better able to withstand thermal stresses than large grained material.

A further description of the present invention will be made in conjunction with the attached sheet of drawings which illustrates one-embodiment of the process.

In the drawings:

FIGURE 1 is a cross-sectional view of a shaping die assembly employed for shaping the preform of the invention and for changing the structure thereof in selected areas;

FIGURE 2 is a view similar to FIG. 1 but showing the elements after the shaping operation has been completed;

FIGURE 3 is a view in elevation of a completed turbine bucket produced according to the present invention;

FIGURE 4 is a view in elevation of a preform produced according to the present invention, said preform having a flaw therein;

FIGURE 5 is a greatly magnified cross sectional view taken substantially along the line VV of FIG. 4 to illustrate more particularly the flaw;

FIGURE 6 is a view similar to FIG. 5 but illustrating the condition of the end of the vane after the flaw therein has been ground down;

FIGURE 7 is a view similar to FIG. 6 but illustrating the condition of the vane after the hot working operatio which eliminates the flaw; and

FIGURE 8 is a graph illustrating the effects of varying amounts of reduction on the stress rupture life of an alloy particularly useful in the process of the invention.

As shown on the drawings:

In FIG. 1, numeral 10 indicates generally a shaping or coining die composed of an upper die member 11 and a lower die member 12 having cooperating shaping surfaces 11a and 12a which cooperate to define the shape of the ultimate article. In the particular embodiment shown in the drawings, the article being manufactured is a turbine bucket of the type shown in FIG. 3, including a relatively massive root portion 13, an arcuate twisted vane portion 1 4 and a shroud 16 located at the upper end of the vane portion 14.

The shaping die 10 is used to hot work a preform 17 which has preferably been produced by vacuum melting, as previously explained. As seen in FIG. 1, the pre form 17 is somewhat oversized in varying degrees at selected portions of the vane portion '14. The dotted line 18 indicates the finished dimensions of the article, and it will be apparent that the greatest amount of excess metal occurs in the vane portion to provide a bulge 17a near the leading edge of the vane, and a bulge 1715 at the trailing edge of the vane.

The preform, in its as-cast structure, consists generally of relatively coarse crystals in the form of dendrites 19 which extend not only through the vane portion 14 but through the root portion 13 and the shroud 16.

The die sections are preferably heated to a temperature on the order of 1950 to 2150 F. for nickel base alloys. As the two die members '11 and 12 are brought together as illustrated in FIG. 2 of the drawings, the shaping surface 11a contacts the bulges 17a and 17b before it contacts the remainder of the vane portion so as to compress those portions and bring them down to the final desired dimensions. Simultaneously, the hot working thus provided to the leading and trailing edges of the vane portion changes the dendritic structure characteristic of those areas to a relatively fine grain structure as illustrated in FIG. 2 of the drawings which includes a large number of relatively finely divided crystals 21 at both the leading and trailing edges of the vane. The central portion of the vane still contains a substantial number of coarse crystals. because the hot working is not nearly as severe. in that portion of the vane as it is in the edges.

Along with the leading and trailing edges of the vane, the tip end of the vane directly below the shroud 16, and indicated in FIG. 3 as reference numeral 14a, is also hot worked to provide the finely divided crystalline structure at that portion of the vane. Thus, the outer marginal edges at the sides and the top of the vane portion 14 are all hot worked to produce a fatigue resistant structure. The inner boundary line of the hot worked area has been generally indicated by the dashed line 23 in FIG. 3.

The flow of metal occurring during the hot working operation also produces a pair of flashes 24 at opposite ends of the vane portion 14, as illustrated in FIG. 2. These flashes 24 can be easily trimmed from the vane portion, after removal from the shaping die.

The process of the present invention lends itself very conveniently to the repair of minor surface defects which may occur from the casting operation. For example, in FIG. 4 there is illustrated a preform 26 having a relatively massive root 27, a vane portion 28 and a shroud 29. The manufacturers specifications usually require that the vane portion 28 be free of surface defects particularly in the area of the leading edge and the trailing edge of the vane. The same requirement exists, but to a lesser degree, in the portion 28a of the vane directly below the shroud, and the portion 28b directly above the root 27.

In FIG. 4 there is illustrated a small surface flaw 31 occurring within the portion of the vane (outlined by the dotted line 32) which must be free of flaws. The flaw 31, in greatly magnified form, is illustrated in FIG. 5 and usually contains some sharp corners which must be eliminated prior to coining or shaping. Accordingly, the flaw 31 is ground down as illustrated in FIG. 6 of the drawings to provide a relatively smoothly contoured depression 32 which is more amenable to the hot working operation.

In FIG. 7, the dotted line 33 represents the original dimensions of the preform and the solid line 34 represents the final, hot worked dimensions.

The hot working operation produces a flash 36 at the edges of the vane portion, and also serves to eliminate the shallow depression 32 by virtue of the flow of metal which occurs during the hot working operation.

The substantial improvement in stress rupture life of an as-cast specimen as compared to a wrought specimen is illustrated in the graph of FIG. 8. After a reduction of about 30%, the stress rupture life of the sample alloy levels 011 at about hours, but in the unreduced or as-cast condition, the stress rupture life of the same alloy was in the neighborhood of 400 hours. These difierences in stress rupture life are directly related to the ability of the alloy to resist creep at elevated temperatures.

From the foregoing, it will be evident that the process of the present invention provides for the manufacture of an integral structure with localized areas particularly adapted to resist those stresses which they encounter in normal operation. It will also be evident that various modifications can be made to the described embodiments without departing from the scope of the present invention.

I claim as my invention:

1. The method of making turbine blades and the like which comprises casting a precipitation hardenable nickel base alloy containing from about 13 to 21% chromium under a reduced pressure of less than one millimeter of mercury, cooling the casting slowly to produce an oversized blade preform having a coarse grained dendritic structure, then hot working the leading and trailing edges of'said preform at temperatures of from 1950 to 2150 F.

to bring said edges down to the desired size and to reduce the grain size in the hot worked portions substantially below that existing in the remaining portions of the blade.

2. The method of making turbine blades and the like which comprises casting a precipitation hardenable nickel base alloy containing from about 13 to 21% chromium and from 10 to 35% cobalt under a reduced pressure of less than one millimeter of mercury, cooling the casting slowly to produce an oversized blade preform having a coarse grained dendritic structure, then hot working the leading and trailing edges of said preform at temperatures of from 1950 to 2150 F. to bring said edges down to the desired size and to reduce the grain size in the hot worked portions substantially below that existing in the remaining portions of the blade.

References Cited in the file of this patent UNITED STATES PATENTS FOREIGN PATENTS Great Britain Nov. 30, 1880 Great Britain Nov. 7, 1951 

